The technical field of this invention is that of nondestructive materials characterization, particularly quantitative, model based characterization of surface, near surface, and bulk material condition for flat and curved parts or components using eddy current sensors. Characterization of bulk material condition includes (1) measurement of changes in material state caused by fatigue damage, creep damage, thermal exposure, or plastic deformation; (2) assessment of residual stresses and applied loads; and (3) assessment of processing related conditions, for example from shot peening, roll burnishing, thermal spray coating, or heat treatment. It also includes measurements characterizing material, such as alloy type, and material states, such as porosity and temperature. Characterization of surface and near surface conditions includes measurements of surface roughness, displacement or changes in relative position, coating thickness, and coating condition. Each of these also includes detection of electromagnetic property changes associated with single or multiple cracks. Spatially periodic field eddy current sensors have been used to measure foil thickness, characterize coatings, and measure porosity, as well as to measure property profiles as a function of depth into a part, as disclosed in U.S. Pat. Nos. 5,015,951 and 5,453,689.
Conventional eddy current sensing involves the excitation of a conducting winding, the primary, with an electric current source of prescribed frequency. This produces a time varying magnetic field at the same frequency, which in turn is detected with a sensing winding, the secondary. The spatial distribution of the magnetic field and the field measured by the secondary is influenced by the proximity and physical properties (electrical conductivity and magnetic permeability) of nearby materials. When the sensor is intentionally placed in close proximity to a test material, the physical properties of the material can be deduced from measurements of the impedance between the primary and secondary windings. Traditionally, scanning of eddy current sensors across the material surface is then used to detect flaws, such as cracks.
For the inspection of structural members in an aircraft, power plant, etc., it is desirable to detect and monitor material damage, crack initiation and crack growth due to fatigue, creep, stress corrosion cracking, etc. in the earliest stages possible in order to verify the integrity of the structure. This is particularly critical for aging aircraft, where military and commercial aircraft are being flown well beyond their original design lives. This requires increased inspection, maintenance, and repair of aircraft components, which also leads to escalating costs. For example, the useful life of the current inventory of aircraft in the U.S. Air Force (e.g., T 38, F 16, C 130E/H, A 10, AC/RC/KC 135, U 2, E 3, B 1B, B 52H) is being extended an additional 25 years at least [Air Force Association, 1997, Committee, 1997]. Similar inspection capability requirements also apply to the lifetime extension of engine components [Goldfine, 1998].
Safely supporting life extension for structures requires both rapid and cost effective inspection capabilities. The necessary inspection capabilities include rapid mapping of fatigue damage and hidden corrosion over wide areas, reduced requirements for calibration and field standards, monitoring of difficult to access locations without disassembly, continuous on line monitoring for crack initiation and growth, detection of cracks beneath multiple layers of material (e.g., second layer crack detection), and earlier detection of cracks beneath fastener heads with fewer false alarms. In general, each inspection capability requires a different sensor configuration.
The use of eddy current sensors for inspection of critical locations is an integral component of the damage tolerance and retirement for cause methods used for commercial and military aircraft. The acceptance and successful implementation of these methods over the last three decades has enabled life extension and safer operation for numerous aircraft. The corresponding accumulation of fatigue damage in critical structural members of these aging aircraft, however, is an increasingly complex and continuing high priority problem. Many components that were originally designed to last the design life of the aircraft without experiencing cracking (i.e., safe life components) are now failing in service, both because aircraft remain in service beyond original design life and, for military aircraft, because expanded mission requirements expose structures to unanticipated loading scenarios. New life extension programs and recommended repair and replacement activities are often excessively burdensome because of limitations in technology available today for fatigue detection and assessment. Managers of the Aircraft Structural Integrity Program (ASIP) are often faced with difficult decisions to either replace components on a fleet wide basis or introduce costly inspection programs.
Furthermore, there is growing evidence that (1) multiple site damage or multiple element damage may compromise fail safety in older aircraft, and (2) significant fatigue damage, with subsequent formation of cracks, may occur at locations not considered critical in original fatigue evaluations. In application of damage tolerance, inspection schedules are often overly conservative because of limitations in fatigue detection capability for early stage damage. Even so, limited inspection reliability has led to numerous commercial and military component failures.
A better understanding of crack initiation and short crack growth behavior also affects both the formulation of damage tolerance methodologies and design modifications on new aircraft and aging aircraft. For safe life components, designed to last the life of the aircraft, no inspection requirements are typically planned for the first design life. Life extension programs have introduced requirements to inspect these “safe life” components in service since they are now operating beyond the original design life. However, there are also numerous examples of components originally designed on a safe life basis that have failed prior to or near their originally specified design life on both military and commercial aircraft.
For safe life components that must now be managed by damage tolerance methods, periodic inspections are generally far more costly than for components originally designed with planned inspections. Often the highest cost is associated with disassembly and surface preparation. Additionally, readiness of the fleet is directly limited by time out of service and reduced mission envelopes as aircraft age and inspection requirements become more burdensome. Furthermore, the later an inspection uncovers fatigue damage the more costly and extensive the repair, or the more likely replacement is required. Thus, inspection of these locations without disassembly and surface preparation is of significant advantage; also, the capability to detect fatigue damage at early stages can provide alternatives for component repair (such as minimal material removal and shotpeening) that will permit life extension at a lower cost than current practice.
In general, fatigue damage in metals progresses through distinct stages. These stages can be characterized as follows [S. Suresh, 1998]: (1) substructural and microstructural changes which cause nucleation of permanent damage, (2) creation of microscopic cracks, (3) growth and coalescence of these microscopic flaws to form ‘dominant’ cracks, (4) stable propagation of the dominant macrocrack, and (5) structural instability or complete fracture.
Although there are differences of opinion within the fatigue analysis community, Suresh defines the third stage as the demarcation between crack initiation and propagation. Thus, the first two of the above stages and at least the initial phase of Stage 3 are generally thought of, from a practical engineering perspective, as the crack initiation phase.
In Stage 1, microplastic strains develop at the surface even at nominal stresses in the elastic range. Plastic deformation is associated with movement of linear defects known as dislocations. In a given load cycle, a microscopic step can form at the surface as a result of localized slip forming a “slip line”. These slip lines appear as parallel lines or bands commonly called “persistent slip bands” (PSBs). Slip band intrusions become stress concentration sites where microcracks can develop.
Historically, X ray diffraction and electrical resistivity are among the few nondestructive methods that have been explored for detection of fatigue damage in the initiation stages. X ray diffraction methods for detection of fatigue damage prior to microcracking have been investigated since the 1930's [Regler, 1937; Regler, 1939]. In these tests, fatigue damage was found to be related to diffraction line broadening. More recently Taira [1966], Kramer [1974] and Weiss and Oshida [1984] have further developed the X ray diffraction method. They proposed a self referencing system for characterization of damage, namely the ratio of dislocation densities as measured 150 micrometers below the surface to that measured 10-50 micrometers below the surface. The data obtained to date suggest that in high strength aluminum alloys the probability of fatigue failure is zero for dislocation density ratios of 0.6 or below. However, it is generally impractical to make such measurements in the field.
Electrical resistivity also provides a potential indication of cumulative fatigue damage. This is supported by theory, since an increase in dislocation density results in an increase in electrical resistivity. Estimates suggest that, in the case of aluminum, depending on the increase in the density of dislocations in the fatigue damage zone, the resistivity in the fatigue affected region may increase by up to 1% prior to formation of microcracks. These estimates are based on dislocation densities in the fatigue damage zone up to between 2(1011 cm2 to 1012 cm 2 and a resistivity factor of 3.3(10 19 ((cm3 [Friedel, 1964].